1. Field of the Invention
The present invention relates to a gas turbine that can guarantee optimal clearance dimensions between rotor blades and a partition ring during operation.
2. Description of the Related Art
FIG. 2 shows an example of the schematic structure of a gas turbine plant. The gas turbine plant shown in this figure comprises a compressor 10, a combustor 20, and a gas turbine 30. In this gas turbine plant, the compressed air that has been compressed by the compressor 10 is supplied to the combustor 20, mixed with fuel supplied separately, and burned. The combusted gas generated by this combustion is supplied to the gas turbine 30, and a rotational drive force is generated by the gas turbine.
Specifically, as shown in FIG. 3, inside the gas turbine 30, a plurality of rotor blades 32 installed on the rotor 31 side and a plurality of stationary blades 33 installed on the stationary side on the periphery of the rotor 31 (not illustrated) are disposed alternating in the axial direction (the left to right direction in the figure) of the rotor 31, and a combustion gas flow path 34 that passes therethrough is formed. Thereby, when the combustion gas supplied into the gas turbine 30 passes through the combustion gas flow path 34, a rotational force is applied to the rotor 31 due to the rotation of each of the rotor blades 32. This rotational force rotates the generators (not illustrated) connected to the rotor 31 to generate electricity.
However, in this gas turbine 30, in order to introduce combustion gas into the interior, the components which have been heated to a high temperature must be cooled, and as shown in FIG. 2, a structure is generally used in which for example, a portion of the compressed air that has been compressed by the compressor 10 is incorporated into an a bleed and used to cool each of the rotor blades 32 and the stator blades 33.
Among these multistage structures, the details of the bleed intake structure in the first stage will be explained below with reference to FIG. 4. Moreover, this figure is an enlargement corresponding to part A in FIG. 3, where the left side of the page is the upstream flow direction of the combustion gas and the right side of the page is the downstream side.
On the outer periphery of each of the rotor blades 32, a partition ring 35 having a ring shape is formed so as to conform to these rotor blades 32, and the partition ring 35 is supported and anchored via the pair of thermally insulating rings 36a and 36b. In order to avoid contact between each of the rotor blades 32 and the partition ring 35, a predetermined clearance c is provided between the outer peripheral edge of each of the rotor blades 32 and the inner peripheral surface of the partition rings 35.
The flow path 38a that opens towards the partition ring 35 is formed by the first stage rotor blades 32, and the bleed f brought in from outside the gas turbine 30 is introduced.
Each of the thermally insulating rings 36a and 36b are a pair of ring shaped parts separated from each other, and in the outside peripheral part thereof, they are separately anchored within the first stage blade ring 38.
In addition, a ring shaped impinging plate 39 and a partition ring 35 are installed and anchored is a state in which they are interposed between the thermally insulating rings 36a and 36b. A plurality of through holes 39a are bored at substantially equal intervals with respect to the outer peripheral surface of a partition ring 35 for distributing and supplying the vapor oil f taken in via the flow path 38a. 
The flanges 35a and 35b are formed at the upstream side and the downstream side of the outer peripheral surface of the partition rings 35, and these flanges 35a and 35b are engaged in a recess formed in each of the thermally insulating rings 36a and 36b. Similarly, both ends of the impinging plate 39 engage in the recesses formed in each of the thermally insulating rings 36a and 36b. 
On the partition rings 35, a plurality of cooling paths 35c that pass from the upstream side of the outer peripheral surface thereof through the interior to the downstream side end surface are formed.
The above explains the first stage structure among the plurality of stages, but the second and subsequent stages positioned on the downstream side therefrom also have substantially the same structure.
However, the clearance c changes due to the differences in thermal expansion between each of the structural components. When this becomes excessively large, there is the problem that the capacity of the gas turbine 30 drastically deteriorates. From this point of view, using an optimal clearance that takes into consideration the differences in thermal expansion between each of the structural components is necessary during the design stage.
However, actually the amount of thermal deformation of each component (for example, the blade rings of each stage starting with the first stage blade ring 38) differs at each of the stages, and thus optimal design is difficult. That is, because the flow conditions (temperature and the like) of the bleed f that cools each of the stages differs for each stage, there is the problem that it is difficult to design with a high precision the clearance c for each of the stages that conforms to the actual shape during operation.
Among these stages, the difference in thermal expansion between the first stage and the second stage is severe, and for example, when the temperature of the members of the second stage blade rings 38A, which are the blade rings of the second stage, is approximately 360xc2x0 C., at the first stage blade ring 38, the temperature of the members is a comparatively high 450xc2x0 C., and thus the clearance c of the first stage has a tendency to become larger than that of the second stage during operation.
The combustion gas flow path 34 has a shape in which the width dimension of the flow path gradually widens from the upstream side to the downstream side at each stage, and thus for the same clearance c, at the upstream first stage, whose flow path width is comparatively narrow, the amount of fluctuation of the clearance c with respect to the flow path width greatly influences the power of the gas turbine 30. Against this background, a structure in which the clearance c is optimal during operation is desired.
In consideration of the above, it is an object of the present invention to provide a gas turbine that can minimize the clearance between each of the rotor blades and the partition rings during operation.